US11420759B2 - Supersonic aircraft and method of reducing sonic booms - Google Patents
Supersonic aircraft and method of reducing sonic booms Download PDFInfo
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- US11420759B2 US11420759B2 US17/044,178 US201917044178A US11420759B2 US 11420759 B2 US11420759 B2 US 11420759B2 US 201917044178 A US201917044178 A US 201917044178A US 11420759 B2 US11420759 B2 US 11420759B2
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- 238000000034 method Methods 0.000 title claims abstract description 27
- 230000005405 multipole Effects 0.000 claims description 22
- 230000002401 inhibitory effect Effects 0.000 claims description 3
- 238000010586 diagram Methods 0.000 description 14
- 230000000694 effects Effects 0.000 description 9
- 230000035939 shock Effects 0.000 description 5
- 238000005516 engineering process Methods 0.000 description 4
- 230000009467 reduction Effects 0.000 description 3
- 230000000052 comparative effect Effects 0.000 description 2
- 230000002349 favourable effect Effects 0.000 description 2
- 230000010006 flight Effects 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 230000032258 transport Effects 0.000 description 2
- 238000004458 analytical method Methods 0.000 description 1
- 230000008901 benefit Effects 0.000 description 1
- 238000004364 calculation method Methods 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 230000002596 correlated effect Effects 0.000 description 1
- 230000000875 corresponding effect Effects 0.000 description 1
- 230000007423 decrease Effects 0.000 description 1
- 239000002360 explosive Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000000644 propagated effect Effects 0.000 description 1
- 230000001629 suppression Effects 0.000 description 1
- 238000010200 validation analysis Methods 0.000 description 1
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 description 1
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Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
- B64C23/04—Influencing air flow over aircraft surfaces, not otherwise provided for by generating shock waves
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C30/00—Supersonic type aircraft
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C5/00—Stabilising surfaces
- B64C5/02—Tailplanes
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/02—Aircraft characterised by the type or position of power plants
- B64D27/16—Aircraft characterised by the type or position of power plants of jet type
- B64D27/20—Aircraft characterised by the type or position of power plants of jet type within, or attached to, fuselages
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D29/00—Power-plant nacelles, fairings, or cowlings
- B64D29/04—Power-plant nacelles, fairings, or cowlings associated with fuselages
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/04—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of exhaust outlets or jet pipes
- B64D33/06—Silencing exhaust or propulsion jets
Definitions
- the present invention relates to a supersonic aircraft such as a supersonic passenger aircraft and a method of reducing sonic booms in such a supersonic aircraft.
- shock waves generated from the respective parts of the airframe integrate as they propagate long distances in the atmosphere, and are observed over land as N-type pressure waveforms that cause two abrupt pressure fluctuations, which are audible to humans as instantaneous explosive sounds. It is generally called “sonic boom”. Concorde, retired in 2003, had not been permitted to fly over land at supersonic speed due to the sonic boom and its flights had been restricted only to flights over water, and the sonic boom is an important technical problem for realizing future supersonic passenger aircraft.
- Non-Patent Literature 1 has disclosed a technology of calculating an airframe equivalent cross-sectional area distribution (total cross-sectional area equivalent to cross-sectional area and lift of the airframe) for reducing sonic booms for aircraft design conditions (airframe length, airframe weight, flight Mach number, flight altitude, and the like).
- a specific airframe shape is designed by using the technology of Non-Patent Literature 1.
- the sonic-boom reduction airframe shape concept proposed by Patent Document 1 has been applied to experimental aircraft and the flight demonstration has been performed.
- Patent Literature 1 and Non-Patent Literatures 1 to 3 all have disclosed technologies for reducing sonic booms by the airframe shape design, but in order to apply them to actual aircraft, it is necessary to also consider the influence of engine exhaust on sonic booms.
- a supersonic aircraft includes: an engine nacelle mounted on a fuselage of an airframe; and a pair of shielding plates that is disposed on the airframe so as to sandwich engine exhaust discharged from a jet engine accommodated in the engine nacelle and inhibits the engine exhaust from wrapping downward around the airframe.
- the pair of shielding plates inhibit the engine exhaust from wrapping downward around the airframe, sonic booms due to the engine exhaust can be reduced.
- the supersonic aircraft according to the embodiment of the present invention may further include a horizontal tail disposed behind the engine nacelle, in which the pair of shielding plates may be disposed on the horizontal tail.
- the pair of shielding plates may be each inclined outward from the airframe.
- the pair of shielding plates may be each inclined at an angle that is determined by using a third-order pole in a multipole method as an index.
- the pair of shielding plates may each include a camber inside the airframe.
- the pair of shielding plates not only reduce the negative pressure but also positively generate the positive pressure to thereby reduce sonic booms.
- this positive pressure can be increased and the effect of reducing sonic booms by the shielding plates can be enhanced.
- the pair of shielding plates may be opposite to each other with an opposing distance that is longer in a direction from a front to a rear of the airframe.
- the positive pressure can be enhanced by the pair of shielding plates with this configuration and the effect of reducing sonic booms by the shielding plates can be enhanced.
- the supersonic aircraft according to the embodiment of the present invention may further include an aft fuselage lifting surface provided behind the engine nacelle, in which the pair of shielding plates is disposed on the aft fuselage lifting surface and has a function as a V tail.
- the pair of shielding plates may each draw inverted Mach cones from positions at which sonic booms are capable of being reduced by providing the pair of shielding plates and may be disposed at positions based on the inverted Mach cone.
- a method of reducing sonic booms is a method of reducing sonic booms of a supersonic aircraft with an engine nacelle mounted on a fuselage of an airframe, the method including: disposing a pair of shielding plates on the airframe so as to sandwich engine exhaust discharged from a jet engine accommodated in the engine nacelle; and inhibiting the engine exhaust from wrapping downward around the airframe by the pair of shielding plates.
- the method of reducing sonic booms according to the embodiment of the present invention may further include drawing inverted Mach cones from positions at which pressure is to be increased by disposing the pair of shielding plates and disposing the pair of shielding plates at positions based on the inverted Mach cones.
- the method of reducing sonic booms according to the embodiment of the present invention may further include: setting a first position on a plane of symmetry that crosses perpendicularly to a center of the airframe, the first position being the position at which the pressure is to be increased by disposing the pair of shielding plates and a second position and a third position shifted by a predetermined angle in first and second directions of a circumferential direction around the fuselage of the airframe from the first position, the second position and the third position being each the position at which the pressure is to be increased by disposing the pair of shielding plates; drawing first to third inverted Mach cones from the first to third positions, respectively; positioning a rear end of a shielding plate of the pair of shielding plates, which is located on a side of the second direction, at a point on a parabola at which the first inverted Mach cone and the second inverted Mach cone intersect; and positioning a rear end of a shielding plate of the pair of shielding plates, which is located on a side of the first direction
- the method of reducing sonic booms according to the embodiment of the present invention may further include defining an angle at which each of the pair of shielding plates is inclined outward from the airframe by using the third-order pole in the multipole method as the index.
- the method of reducing sonic booms according to the embodiment of the present invention may further include defining an angle at which each of the pair of shielding plates is inclined outward from the airframe by using a difference between a correction amount of a pressure distribution according to the multipole method immediately below the airframe in a case where the pair of shielding plates is provided and a correction amount of a pressure distribution according to the multipole method immediately below the airframe in a case where the pair of shielding plates is not provided or a difference between a third-order pole distribution in a case where the pair of shielding plates is provided and a third-order pole distribution in a case where the pair of shielding plates is not provided as an index.
- sonic booms due to engine exhaust can be reduced.
- FIG. 1 A plan view showing the outer appearance of a supersonic aircraft according to an embodiment of the present invention.
- FIG. 2 A side view showing the outer appearance of the supersonic aircraft shown in FIG. 1 .
- FIG. 3 A front view showing the outer appearance of the supersonic aircraft shown in FIG. 1 .
- FIG. 4 A diagram showing a definition of a coordinate axis, which is required for describing the embodiment of the present invention.
- FIG. 5 A diagram showing a definition of an angle in a circumferential direction, which is required for describing the embodiment of the present invention.
- FIG. 6 A diagram showing a definition of a position shifted by a predetermined angle in the circumferential direction, which is required for describing the embodiment of the present invention.
- FIG. 9 A diagram showing a state in which the inverted Mach cone is drawn from each position when determining the position of the fin according to the embodiment of the present invention by the inverted Mach cone.
- FIG. 10 A diagram showing a definition of an angle of the fin according to the embodiment of the present invention.
- FIG. 11 A graph showing a relationship between a pole order in a multipole method and a correlation coefficient.
- FIG. 12 A graph showing an example of a waveform (waveform A) of a difference in correction amount of a pressure distribution between a case where the fins are provided and a case where the fins are not provided and a waveform (waveform B) of a difference in third-order pole distribution in the multipole method between the case where the fins are provided and the case where the fins are not provided according to the embodiment of the present invention.
- FIG. 13 A side view of the fin according to the embodiment of the present invention.
- FIG. 14 A diagram showing a shape of the A-A cross-section of FIG. 13 .
- FIG. 15 A graph showing a pressure distribution in a near field immediately below an airframe in a case where the fins are provided and in a case where the fins are not provided for verifying the effects of the fins according to the embodiment of the present invention.
- FIG. 16 A graph showing a sonic-boom waveform over land immediately below the airframe in the case where the fins are provided and in the case where the fins are not provided for verifying the effects of the fins according to the embodiment of the present invention.
- FIG. 17 A diagram showing a pressure-coefficient distribution of a cross-section orthogonal to the X-axis at a predetermined position on the X-axis according to the embodiment of the present invention.
- FIG. 18 A diagram showing a pressure-coefficient distribution of a cross-section orthogonal to the X-axis at a predetermined position on the X-axis in the case where the fins are not provided as a comparative example of FIG. 17 .
- FIG. 19 A diagram showing a pressure-coefficient distribution on a plane of symmetry according to the embodiment of the present invention.
- FIG. 20 A diagram showing a pressure-coefficient distribution on a plane of symmetry in the case where the fins are provided as a comparative example of FIG. 19 .
- FIG. 21 A plan view showing the outer appearance of a supersonic aircraft according to another embodiment of the present invention.
- FIG. 22 A side view showing the outer appearance of the supersonic aircraft shown in FIG. 21 .
- FIG. 23 A front view showing the outer appearance of the supersonic aircraft shown in FIG. 21 .
- FIG. 1 is a plan view showing the outer appearance of a supersonic aircraft according to the embodiment of the present invention.
- FIG. 2 is a side view thereof and
- FIG. 3 is a front view thereof.
- the supersonic aircraft includes a pair of engine nacelles 12 R, 12 L mounted on a fuselage 11 of an airframe 10 , fins 13 R, 13 L as a pair of shielding plates that inhibits engine exhaust 15 discharged from jet engines (not shown) accommodated in the engine nacelles 12 R, 12 L from wrapping downward around the airframe 10 , and a pair of horizontal tails 14 R, 14 L disposed behind the engine nacelles 12 R, 12 L.
- the fins 13 R, 13 L are typically disposed on the horizontal tails 14 R, 14 L of the airframe 10 so as to sandwich the engine exhaust 15 , respectively.
- the pair of fins 13 R, 13 L and the pair of horizontal tails 14 R, 14 L are respectively arranged in plane symmetry with respect to a plane of symmetry 16 that crosses perpendicularly to the axis of the airframe 10 .
- the fin 13 R is mounted on the horizontal tail 14 R and the fin 13 L is mounted on the horizontal tail 14 L.
- FIG. 4 is a diagram showing a definition of coordinate axes required for describing this embodiment.
- the length of the fuselage 11 is 1.
- the Mach number was set to 1.6 at the cruising speed of the supersonic aircraft.
- positions of the fins 13 R, 13 L are determined on the basis of inverted Mach cones.
- FIG. 5 is a diagram showing a definition of an angle in a circumferential direction, which is required for describing this embodiment.
- FIG. 6 is a diagram showing a definition of a point shifted by a predetermined angle in the circumferential direction, which is required for describing this embodiment.
- a near field position to increase the pressure is determined from the viewpoint of sonic-boom reduction.
- a pressure waveform on an X coordinate axis on the plane of symmetry is shown in the graph of FIG. 7 and a pressure waveform on an X coordinate axis on the plane 16 B after the plane of symmetry 16 is rotated by 30 degrees is shown in the graph of FIG. 8 .
- a waveform 70 d represents a case where the fins are provided and a waveform 70 e represents a case where the fins are not provided.
- p on the Y-axis is a uniform flow static pressure and dp is a difference from the uniform flow static pressure.
- the inverted Mach cones 20 A, 20 B are drawn from two points of the point A and the point B, respectively, and the position of the fin 13 L is set by using a point on a parabola at which these intersect as a position of a wing tip trailing edge of the fin 13 L (Step 2 ).
- Step 3 a viewpoint of whether the position of the fin 13 L set in Step 2 is a favorable position for shielding the exhaust jet is added, in other words, the point on the parabola or the points A and B is corrected in a manner that depends on needs and Steps 1 and 2 described above are repeated (Step 3 ).
- the viewpoint of whether the position of the fin 13 L set in Step 2 is the favorable position for shielding the exhaust jet is determined on the basis of a consideration that a significant shielding effect is not provided if the position obtained from the inverted Mach cone is far above the exhaust jet, for example.
- FIG. 10 is a diagram showing a definition of an angle of the fin according to this embodiment.
- the pair of fins 13 R, 13 L is each inclined outward from the airframe. This inclination angle is defined as the angle of the fin.
- the X-axis is separated from the airframe 1 by a distance that is 0.3 times as long as the fuselage length.
- the pressure waveform changes as this position moves away from the airframe 1 (see FIGS. 19 and 20 to be described later).
- the pressure distribution on the X-axis is defined as a reference waveform.
- a virtual pressure waveform obtained by correcting the reference waveform in consideration of the three-dimensional circumferential pressure propagation is defined as a multipole waveform. More specifically, correcting the reference waveform refers to replacing the airframe with an equivalent multipole distribution and virtually adding a pressure waveform deformation due to the pressure propagation in the circumferential direction to the reference waveform while reflecting the intensity decay of the pressure propagation in the circumferential direction given for each pole order.
- the multipole waveform ⁇ the reference waveform is defined as a correction amount.
- the fin inclination angle is set by using this correction amount as an index.
- FIG. 11 is a graph showing a relationship between a pole order and a correlation coefficient regarding a correlation between a difference relating to the correction amount between the case where the fins are provided and the case where fins are not provided and a difference relating to the pole distribution in the multipole method between the case where the fins are provided and the case where fins are not provided.
- the correlation coefficient of the third-order pole is large. That is, the third-order pole can best express the pressure shielding by the fins 13 R, 13 L. Therefore, it can be said that it is effective to set the fin inclination angle by using the third-order pole distribution as an index. More specifically, this third-order pole distribution is an index representing the strength of the up-and-down asymmetry of the flow field. This third-order pole index is defined as a multipole correlation.
- waveform A of the difference in correction amount between the case where the fins are provided and the case where fins are not provided
- waveform B waveform of the difference in multipole correlation between the case where the fins are provided and the case where fins are not provided
- the waveform A indicates that regarding the peak between 0.95 and 1 of the difference in correction amount, the correction amount at this position is larger in the case where the fins are provided than in the case where the fins are not provided.
- the expansion wave generated by the engine exhaust three-dimensionally wraps in the circumferential direction, such that the correction amount becomes negative. With the fins, this wrapping is suppressed, such that a peak is formed because the correction amount on the negative side decreases. Since this suppression of the wrapping of the expansion wave is effective for reducing sonic booms, it is sufficient to set the fin inclination angle such that this peak is higher.
- the fin inclination angle may be set by using the correction amount as an index, the fin inclination angle may be set by using the multipole correlation as an index, or the fin inclination angle may be set by using both the correction amount and the multipole correlation as indices.
- FIG. 13 is a side view of the fin 13 R and FIG. 14 is an A-A cross-sectional view of the fin 13 R shown in FIG. 13 . It should be noted that the fin 13 L is plane-symmetrical with the fin 13 R with respect to the plane of symmetry 16 .
- each of the fins 13 R, 13 L includes a camber 21 directed toward the inside of the airframe 10 .
- an opposing distance C between the fin 13 R and the fin 13 L is longer in a direction from the front to the rear of the airframe 10 .
- the effect of the sonic-boom reduction can be enhanced by setting the fin shapes of the fins 13 R, 13 L and the opposing distance C between the fins 13 R, 13 L by using the multipole method as in the setting of the fin inclination angle.
- the supersonic aircraft according to this embodiment includes the fins 13 R, 13 L, the influence of the engine exhaust 15 on sonic booms can be reduced.
- FIG. 15 is a graph showing a pressure waveform in a near field immediately below the airframe in a case where the fins 13 R, 13 L are provided ( 70 d ) and in the case where the fins 13 R, 13 L are not provided ( 70 e ).
- FIG. 16 is a graph showing a sonic-boom waveform over land immediately below the airframe in a case where the fins 13 R, 13 L are provided ( 70 d ) and in the case where the fins 13 R, 13 L are not provided ( 70 e ).
- FIGS. 17 and 18 each show a pressure-coefficient distribution on a cross-section orthogonal to the X-axis at a predetermined position (near fin position) on the X-axis.
- FIG. 17 shows the case where the fins 13 R, 13 L are provided and
- FIG. 18 shows the case where the fins are not provided.
- FIGS. 19 and 20 each show a pressure-coefficient distribution on the plane of symmetry 16 .
- FIG. 19 shows the case where the fins 13 R, 13 L are provided and
- FIG. 20 shows the case where the fins 13 R, 13 L are not provided.
- the present invention can be provided not only in the supersonic aircraft having the configuration shown in the above-mentioned embodiment, but also in various forms of supersonic aircraft.
- FIG. 21 is a plan view showing the outer appearance of a supersonic aircraft according to a second embodiment.
- FIG. 22 is a side view thereof and
- FIG. 23 is a front view thereof.
- the supersonic aircraft includes a pair of engine nacelles 112 R and 112 L mounted on a fuselage 111 of an airframe 110 and a pair of shielding plates 113 R, 113 L disposed behind the engine nacelles 112 R, 112 L.
- the pair of shielding plates 113 R, 113 L is mounted on an aft fuselage lifting surface 114 provided behind the engine nacelles 112 R, 112 L so as to be inclined outward from the airframe.
- the pair of shielding plates 113 R, 113 L prevents inhibits the engine exhaust 15 discharged from jet engines (not shown) accommodated in the engine nacelles 112 R, 112 L from wrapping around the fuselage 111 and has a function as a V tail.
- sonic booms can be reduced when the pair of shielding plates 113 R, 113 L is provided as described in the first embodiment.
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- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
Description
- Patent Literature 1: Japanese Patent No. 5057374
- Non-Patent Literature 1: Christine M. Darden: Sonic-Boom Minimization With Nose-Bluntness Relaxation, NASA TP-1348, 1979.
- Non-Patent Literature 2: Todd E. Magee, Peter A. Wilcox, Spencer R. Fugal, and Kurt E. Acheson, Eric E. Adamson, Alicia L. Bidwell, Stephen G. Shaw: System-Level Experimental Validations for Supersonic Commercial Transport Aircraft Entering Service in the 2018-2020 Time Period Phase I Final Report, NASA/CR-2013-217797, 2013.
- Non-Patent Literature 3: John Morgenstern, Nicole Norstrud, Jack Sokhey, Steve Martens, and Juan J. Alonso: Advanced Concept Studies for Supersonic Commercial Transports Entering Service in the 2018 to 2020 Period Phase IFinal Report, NASA/CR-2013-217820, 2013.
β×r=sqrt(Mach2−1)×r=sqrt(1.62−1)×0.3=0.375.
- 1 airframe
- 10 airframe
- 11 fuselage
- 12L engine nacelle
- 12R engine nacelle
- 13L fin
- 13R fin
- 14L horizontal tail
- 14R horizontal tail
- 15 engine exhaust
- 16 plane of symmetry
- 16B surface
- 16B′ surface
- 17 mach line
- 18 nose tip
- 20A inverted Mach cone
- 20B inverted Mach cone
- 20B′ inverted Mach cone
- 21 camber
- 70 d a waveform where fins are provided
- 70 e a waveform where fins are not provided
- 110 airframe
- 111 fuselage
- 112L engine nacelle
- 112R engine nacelle
- 113L shielding plate
- 113R shielding plate
- 114 aft fuselage lifting surface
- C opposing distance
- ϕ angle
Claims (2)
Applications Claiming Priority (4)
Application Number | Priority Date | Filing Date | Title |
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JP2018-073784 | 2018-04-06 | ||
JPJP2018-073784 | 2018-04-06 | ||
JP2018073784A JP7103631B2 (en) | 2018-04-06 | 2018-04-06 | How to reduce supersonic aircraft and sonic booms |
PCT/JP2019/012464 WO2019194002A1 (en) | 2018-04-06 | 2019-03-25 | Supersonic airplane and method for reducing sonic booms |
Publications (2)
Publication Number | Publication Date |
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US20210031935A1 US20210031935A1 (en) | 2021-02-04 |
US11420759B2 true US11420759B2 (en) | 2022-08-23 |
Family
ID=68100367
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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US17/044,178 Active 2039-04-02 US11420759B2 (en) | 2018-04-06 | 2019-03-25 | Supersonic aircraft and method of reducing sonic booms |
Country Status (4)
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US (1) | US11420759B2 (en) |
EP (1) | EP3778387A4 (en) |
JP (1) | JP7103631B2 (en) |
WO (1) | WO2019194002A1 (en) |
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KR101909912B1 (en) * | 2016-08-16 | 2018-10-19 | (주)광주요 | Gypsum mold for porcelain molding and manufacturing method of porcelain using the mold |
US20220268236A1 (en) * | 2021-02-24 | 2022-08-25 | Japan Aerospace Exploration Agency | Supersonic aircraft and method of reducing sonic booms and jet noise |
CN115320879B (en) * | 2022-10-14 | 2022-12-09 | 中国空气动力研究与发展中心低速空气动力研究所 | Method for designing coanda profile of ring control airfoil trailing edge |
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WO2016203015A1 (en) | 2015-06-19 | 2016-12-22 | Centre National De La Recherche Scientifique | System for reducing the installation noise of an aeroplane wing |
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2018
- 2018-04-06 JP JP2018073784A patent/JP7103631B2/en active Active
-
2019
- 2019-03-25 WO PCT/JP2019/012464 patent/WO2019194002A1/en unknown
- 2019-03-25 EP EP19780993.2A patent/EP3778387A4/en active Pending
- 2019-03-25 US US17/044,178 patent/US11420759B2/en active Active
Patent Citations (16)
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US3936017A (en) * | 1973-07-30 | 1976-02-03 | Hawker Siddeley Aviation Limited | Combined noise shield and thrust reverser |
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JP7103631B2 (en) | 2022-07-20 |
US20210031935A1 (en) | 2021-02-04 |
JP2019182125A (en) | 2019-10-24 |
EP3778387A4 (en) | 2021-12-08 |
EP3778387A1 (en) | 2021-02-17 |
WO2019194002A1 (en) | 2019-10-10 |
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